1. Field of the Invention
The present invention relates generally to gas turbine engine, and more specifically to an air cooled blade outer air seal (BOAS) with teeth for an industrial gas turbine engine.
2. Description of the Related Art including Information Disclosed under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
A row or stage of turbine rotor blades rotate within an annular arrangement of ring segments in which blade tips form a small gap with an inner or hot surface of each ring segment. The size of the gap changes due to different thermal properties of the blade and the BOAS or ring segments from a cold sate to a hot state of the turbine. The smaller the gap, the less hot gas leakage will flow between the blade tips and the ring segments.
An IGT engine operates for long periods of time at steady state conditions, as opposed to an aero gas turbine engine that operates for only a few hours before shutting down. Thus, the parts in the IGT engine must be designed for normal operation for these long periods, such as up to 40,000 hours of operation at steady state conditions.
High temperature turbine blade tip shroud heat load is a function of the blade tip section leakage flow. A high leakage flow will induce a high heat load on the blade tip shroud. Therefore, blade tip shroud cooling and sealing issues must be considered as a single problem.
FIGS. 1 through 3 show a prior art blade tip shroud design with a grooved turbine blade tip shroud 31 that includes multiple grooves 32 at angles of 90 to 130 degrees relative to the shroud backing structure, and where the grooves 32 extend into the hot gas flow path fro an entire axial length of the blade outer air seal. The tip shroud 31 is secured to a blade ring carrier 11 through a second piece 13 and forms a cooling air supply cavity 12. An impingement plate 21 with impingement holes 22 directs impingement cooling air onto the backside surface of the tip shroud to produce impingement cooling. As seen in FIG. 2, the tip shroud grooves 32 extend from one end to the opposite end in a straight line with the arrows representing the tip leakage flow direction. FIG. 1 shows the hot gas flow along an outer endwall of an adjacent stator vane assembly 15 and the leakage flow in the blade tip region and in the gap formed between the blade tip and the blade tip shroud 31. FIG. 3 shows the grooves 32 forming teeth 34 with the leakage flow recirculation within the grooves 32 and the leakage flow through the gap.
The main purpose of using a grooved tip shroud in a blade design is to reduce the blade tip leakage and to provide for rubbing capability of the blade tips. The grooved blade tip shroud 31 in FIG. 1 is not cooled. Since the turbine inlet temperature has been steadily increasing over the past few years, cooling of the grooved blade tip shroud becomes necessary.